The present invention relates broadly to a redundant flight control system, and in particular to an in-line actuator monitoring and control apparatus.
In the prior art, it is well known in the field of aircraft control that hydraulic powered systems are generally utilized to position or activate various controlled surfaces. Usually the controlled system is operated by input command signals from the pilot to maintain a particular course, attitude, altitude, etc. However, as the size and speed of modern aircraft increased, so did the requirement that probability of failure in these hydraulic systems be minimized. Thus, it is now common in large and high speed aircraft to provide redundant control systems set with multiple control channels such that failure of one channel does not cause a failure of the entire system, and consequent loss of the aircraft. As a result of these factors, the acceptable response time for switching out one failed channel of a redundant control system is highly critical and therefore, must be minimal.
There have been numerous techniques and concepts utilized to implement redundant control systems. These systems have included standby channels with switch over from one channel to the standby channel upon failure of a single channel, multiple independent control surfaces, each with its own actuator so that failure of one surface does not result in failure of the system, displacement summing channels in which the resulting displacement of the control surface is effected by multiple actuators, each one of which is displaced some fraction of the commanded displacement, and force summed control systems in which the force imparted to the control element by each one of multiple actuators is some fraction of a commanded force. In each case it is important to detect and warn of any failure so that corrective action can be taken before initiation of a chain of events ending in catastrophy. The choice of system depends upon the overall characteristics of the systems such as the tolerable transient shift which may occur upon failure of a channel, size and weight characteristics of the system, and the acceptable probability of failure. In a redundant fly by wire flight control system, when but two channels are operative and a failure occurs in one of the two remaining control surfaces secondary actuators, a requirement exists for a means to detect and indicate which of the two actuators has failed if the system is to continue to operate properly. The present invention provides such a system.